U.S. patent number 6,823,578 [Application Number 09/987,612] was granted by the patent office on 2004-11-30 for one-piece closed-shape structure and method of forming same.
This patent grant is currently assigned to Toyota Motor Sales, U.S.A., Inc.. Invention is credited to Alan H. Anderson, Kathlene K. Bowman, Paul D. Teufel.
United States Patent |
6,823,578 |
Anderson , et al. |
November 30, 2004 |
One-piece closed-shape structure and method of forming same
Abstract
The present invention relates to a one-piece closed-shape
structure and a method for manufacturing a one-piece closed-shape
structure. In particular, the present invention relates to a
one-piece fuselage and a method for manufacturing a one-piece
fuselage. One embodiment of the method of the invention involves
the use of molding technology, tooling technology, the integration
of the molding and tooling technology, and fiber placement to
manufacture a one-piece closed shape structure.
Inventors: |
Anderson; Alan H. (Placentia,
CA), Bowman; Kathlene K. (Lancaster, CA), Teufel; Paul
D. (Los Alamitos, CA) |
Assignee: |
Toyota Motor Sales, U.S.A.,
Inc. (Torrance, CA)
|
Family
ID: |
22938071 |
Appl.
No.: |
09/987,612 |
Filed: |
November 15, 2001 |
Current U.S.
Class: |
264/512 |
Current CPC
Class: |
B29C
33/505 (20130101); B29C 70/342 (20130101); B64C
1/061 (20130101); B29C 70/446 (20130101); B64C
1/064 (20130101); B64C 2211/00 (20130101); B64D
2011/0046 (20130101); Y02T 50/43 (20130101); Y02T
50/40 (20130101); B64C 2001/0072 (20130101); B29L
2031/3082 (20130101) |
Current International
Class: |
B64C
1/00 (20060101); H02K 012/02 () |
Field of
Search: |
;29/569,598,608
;156/173,169 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
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2 424 470 |
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2 664 529 |
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PCT/US98/01740 |
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Other References
John Berry, International Search Report for International
application No. PCT/US01/43091, (Jul. 17, 2002). .
D.V. Rosato et al., "Filament Winding: Its Development,
Manufacture, Applications, and Design", John Wiley & Sons, Inc.
(1964). .
A. Van Wallene, International Search Report for International
application No. PCT/US 01/45750 (Sep. 19, 2002). .
J. Carre, International Search Report for International application
No. PCT/US 02/05094, (Aug. 2, 2002)..
|
Primary Examiner: Dinh; Tien
Attorney, Agent or Firm: Finnegan, Henderson, Farabow,
Garrett & Dunner, L.L.P.
Parent Case Text
I. CROSS-REFERENCE TO RELATED APPLICATION
This application claims the benefit of U.S. Provisional Application
No. 60/248,190, filed Nov. 15, 2000 by Alan H. Anderson, Kathlene
K. Bowman, and Paul D. Teufel and titled ONE-PIECE CLOSED-SHAPE
STRUCTURE AND METHOD OF FORMING SAME, the disclosure of which is
expressly incorporated herein by reference.
Claims
What is claimed is:
1. A method of manufacturing a one-piece closed-shape structure
using a mandrel, comprising: preparing the mandrel, wherein the
mandrel comprises a bag and an armature; applying a frame mandrel
to the mandrel to form a frame for the structure; filling the
mandrel and the frame mandrel with media; applying a curable resign
to a fiber; applying the fiber over the mandrel and frame mandrel
to form the structure; curing the structure; removing the media
from the mandrel and frame mandrel; and extracting the mandrel and
frame mandrel from the structure.
2. The method of claim 1, wherein preparing further comprises:
placing the armature through the bag; and conforming the shape of
the bag to a desired shape of the structure.
3. The method of claim 2, wherein conforming further comprises:
sealing the bag; placing the armature and the bag in a form tool;
and conforming the shape of the bag to the form tool.
4. The method of claim 3, wherein conforming further comprises:
filling a space between the armature and the bag with air; and
creating a vacuum between the form tool and the bag to force the
bag to conform to the shape of the form tool.
5. The method of claim 1, wherein applying a frame mandrel further
comprises: applying a frame ply to an exterior of the bag; and
applying the frame mandrel over the frame ply.
6. The method of claim 1, wherein filling further comprises
compacting the media.
7. The method of claim 6, wherein compacting further comprises
vibrating the mandrel and frame mandrel to aid compaction.
8. The method of claim 1, wherein applying the fiber comprises:
winding the fiber over the mandrel and frame mandrel to form the
structure.
9. The method of claim 8, wherein winding further comprises:
placing a first winding aid on the bag; winding the fiber over the
first winding aid, the frame mandrel, and the mandrel to form an
inner skin; cutting the inner skin to remove the first winding
aids; placing a second winding aid on the inner skin; winding the
fiber over the second winding aid and inner skin to form an outer
skin; and cutting the outer skin to remove the second winding
aids.
10. The method of claim 9, wherein placing second winding aids
further comprises placing a core piece on the inner skin.
11. The method of claim 1, wherein curing further comprises:
placing a mold around an exterior of the structure; sealing the
mold; placing the mold in a heating device; and applying heat to
the mold using the heating device.
12. The method of claim 11, wherein curing further comprises:
creating a vacuum in the mandrel; and creating a vacuum in the
frame mandrel.
13. The method of claim 1, wherein curing further comprises:
placing a mold around an exterior of the structure; sealing the
mold; placing the mold in an autoclave; and applying pressure to
the mold using the autoclave.
14. The method of claim 1, wherein the structure is a fuselage of
an aircraft.
15. A computer-implemented method of manufacturing a one-piece
closed-shape structure using a mandrel, comprising: preparing the
mandrel, wherein the mandrel comprises a bag and an armature;
applying a frame mandrel to the mandrel to form a frame for the
structure; filling the mandrel and the frame mandrel with media;
applying a curable resign to a fiber; applying the fiber over the
mandrel and frame mandrel to form the structure; curing the
structure; removing the media from the mandrel and frame mandrel;
and extracting the mandrel and frame mandrel from the
structure.
16. A method of manufacturing a one-piece closed-shape structure,
using a mandrel comprising: preparing the mandrel, wherein the
mandrel comprises a bag and an armature; placing the armature
through the bag; conforming the shape of the bag to a desired shape
of the structure; applying a frame mandrel to the mandrel to form a
frame of the structure; filling the mandrel and the frame mandrel
with media; applying a curable resign to a fiber; applying the
fiber over the frame mandrel and the bag to form an inner skin;
placing a core piece on the inner skin; applying the fiber over the
core piece and inner skin to form an outer skin; placing a mold
around an exterior of the structure; curing the structure in the
mold; removing the mold from the structure; removing the media from
the mandrel and the mandrel frame; extracting the armature from the
bag; and extracting the bag from the structure.
17. The method of claim 16, wherein conforming further comprises:
sealing the bag; placing the armature and the bag in a form tool;
and conforming the shape of the bag to the form tool.
18. The method of claim 17, wherein conforming further comprises:
filling a space between the armature and the bag with air; and
creating a vacuum between the form tool and the bag to force the
bag to conform to the shape of the form tool.
19. The method of claim 16, wherein applying a frame mandrel
further comprises: applying a frame ply to an exterior of the bag;
and applying a frame mandrel over the frame ply.
20. The method of claim 16, wherein filling further comprises
compacting the media.
21. The method of claim 20, wherein compacting further comprises
vibrating the mandrel and frame mandrel to aid compaction.
22. The method of claim 16, wherein applying the fiber over the
frame mandrel and the bag to form an inner skin comprises: winding
the fiber over the frame mandrel and the bag to form the inner
skin.
23. The method of claim 22, wherein winding further comprises:
placing a winding aid on the bag; winding the fiber over the frame
mandrels, the winding aid, and the bag to form the inner skin; and
cutting the inner skin to remove the winding aid.
24. The method of claim 16, wherein applying the fiber over the
core piece and inner skin to form an outer skin comprises: winding
the fiber over the core piece and inner skin to form the outer
skin.
25. The method of claim 24, wherein winding further comprises:
placing a winding aid on the inner skin; winding the fiber over the
core piece, the winding aid, and the inner skin to form an outer
skin; and cutting the outer skin to remove the winding aid.
26. The method of claim 16, wherein curing further comprises:
sealing the mold; placing the mold in a heating device; and
applying heat to the mold using the heating device.
27. The method of claim 26, wherein curing further comprises:
creating a vacuum in the mandrel; and creating a vacuum in the
frame mandrel.
28. The method of claim 16, wherein curing further comprises:
sealing the mold; placing the mold in an autoclave; and applying
pressure to the mold using the autoclave.
29. The method of claim 16, wherein the one-piece closed-shape
structure is an airplane fuselage.
30. A computer-implemented method of manufacturing a one-piece
closed-shape structure, using a mandrel comprising: preparing the
mandrel, wherein the mandrel comprises a bag and an armature;
placing the armature through the bag; conforming the shape of the
bag to a desired shape of the structure; applying a frame mandrel
to the mandrel to form a frame of the structure; filling the
mandrel and the frame mandrel with media; applying a curable resign
to a fiber; applying the fiber over the frame mandrel and the bag
to form an inner skin; placing a core piece on the inner skin;
applying the fiber over the core piece and inner skin to form an
outer skin; placing a mold around an exterior of the structure;
curing the structure in the mold; removing the mold from the
structure; removing the media from the mandrel and the mandrel
frame; extracting the armature from the bag; and extracting the bag
from the structure.
Description
II. BACKGROUND OF THE INVENTION
A. Field of the Invention
The present invention relates to a one-piece closed-shape structure
and a method for manufacturing a one-piece closed-shape structure.
In particular, the present invention relates to a one-piece
fuselage and a method for manufacturing a one-piece fuselage.
B. Background of the Invention
Since the 1940's and 1950's, aircraft have been manufactured from
lightweight metals, primarily aluminum. More recently, composite
materials (such as fiber reinforced plastics) have been used to
manufacture some aircraft. The manufacture of such aircraft include
the manufacture of the fuselage (the central body of the aircraft),
the wings, and the various other components of the aircraft.
In the manufacture of an aircraft fuselage with metals or
composites, the typical manufacturing process involves the
combination of several pieces that are individually manufactured
and then bonded together to form the fuselage. These multiple steps
have many disadvantages, including both high cost and significant
time.
The creation of a single-piece fuselage would provide many
advantages over fuselages manufactured from the combination of
multiple parts. These advantages potentially include lower cost,
lighter weight, improved integration, safety, improved performance,
noise reduction, improved aerodynamics, and styling
flexibility.
As for lower cost, a one-piece fuselage is less costly to
fabricate, because there is only one part to manufacture, and there
are no fasteners. Thus, the one-piece design saves money in both
the fabrication stage and in combination stage. In addition, the
work areas needed at a manufacturing facility are less for a
one-piece design, because multiple parts require dramatically more
workspace areas.
As for lighter weight, because there are fewer parts to a one-piece
fuselage, and because there are fewer fasteners, a one-piece
fuselage is lighter than a fuselage created from multiple parts.
The lighter the aircraft, the more carrying capacity that the
aircraft will have, which is a substantial benefit.
As for improved integration, a one-piece fuselage is easier to
integrate with the other components of the aircraft, such as the
tail cone, the wings, and the other parts of the aircraft.
Additionally, the interior of a one-piece fuselage would also be
easier to integrate, because there is only one form that must be
properly fitted. Moreover, problems with integration of multiple
parts (such as dimension variation and other fabrication problems)
would be completely eliminated in a one-piece design.
As for safety, a one-piece fuselage offers structural advantages
over a fuselage fabricated from multiple parts. In the initial
fabrication of the one-piece fuselage, the structure may be
designed with safety improvements (such as strengthened areas,
etc.). Additionally, because the one-piece fuselage does not have
most of the fasteners necessary for combining the multiple parts,
the one-piece design is more structurally sound, which provides
increased passenger safety. Also, a one-piece fuselage is more
crashworthy. A one-piece fuselage provides the advantages of an
integrated structure, which has numerous crashworthiness
benefits.
As for improved performance, there are both objective and
subjective improvements. For objective improvements, there is of
course the improved aerodynamics, which results in greater speed.
For subjective improvement, there is the noise reduction, which
results in a more comfortable ride. In some way, all of the
advantages of the one-piece fuselage play a role in improved
performance.
As for noise reduction, because a one-piece fuselage would result
in improved aerodynamics, a further benefit would be a diminution
of air disruption, which results in noise reduction. Any increase
in the smoothness of an aircraft has the benefit of noise
reduction. Thus, to the extent that the creation of a one-piece
fuselage results in the improvement of aerodynamics, there is a
reciprocal decrease in noise.
As for improved aerodynamics, a one-piece fuselage inherently is
more aerodynamic than a fuselage created from the combination of
multiple parts. This improvement in aerodynamics would result from
the absence of seams or joints as well as the absence of rivets or
other external fasteners. In modern aircraft, seams and joints
between the combined parts increase drag and thus diminish
aerodynamics. By omitting the seams and joints in a one-piece
fuselage, aerodynamics would be improved. Also, in modern aircraft,
the external fasteners for flanges and other structure internal to
the fuselage also increase drag and diminish aerodynamics. A
one-piece fuselage would omit most fasteners and would thus improve
aerodynamics.
As for styling flexibility, the capability to create a one-piece
fuselage would provide more opportunities for aircraft design.
Because multiple parts are not combined to create the fuselage,
unique shapes may be possible, that were previously difficult to
achieve. By improving the design and styling of the aircraft with a
one-piece fuselage, it would thus be possible to create a more
attractive aircraft for the market.
Therefore, it is desirable to provide a one-piece fuselage.
For a one-piece fuselage, either metal or composite materials may
be used. Metal has more disadvantages, due to the inability to
fabricate all components of the fuselage in a single step.
Composite materials are thus more advantageous for the fabrication
of a one-piece fuselage, because composite materials may be
fabricated simultaneously.
Therefore, it is further desirable to provide a one-piece fuselage
manufactured from composite materials.
Methods and structures in accordance with the invention provide for
a one-piece structure manufactured from composite materials,
including a one-piece fuselage. One embodiment includes
manufacturing a one-piece fuselage by filament winding. Other
embodiments for manufacturing a one-piece fuselage may also be
used.
III. SUMMARY OF THE INVENTION
Methods and structures consistent with the present invention may
overcome the shortcomings of conventional systems by providing a
one-piece closed shape structure manufactured by composite
materials. Additional objects and advantages of the invention will
be set forth in part in the description, which follows, and in part
will be obvious from the description, or may be learned by practice
of the invention. The objects of the invention will be realized and
attained by means of the elements and combination particularly
pointed out in the appended claims.
In accordance with an embodiment of the present invention, a method
of manufacturing a one-piece closed-shape structure using a mandrel
comprises: preparing the mandrel, wherein the mandrel comprises a
bag and an armature; applying a frame mandrel to the mandrel to
form a frame for the structure; filling the mandrel and the frame
mandrel with media; applying a curable resign to a fiber; applying
the fiber over the mandrel and frame mandrel to form the structure;
curing the structure; removing the media from the mandrel and frame
mandrel; and extracting the mandrel and frame mandrel from the
structure.
In accordance with another embodiment of the present invention,
preparing further comprises: placing the armature through the bag
and conforming the shape of the bag to a desired shape of the
structure. This embodiment may also include sealing the bag;
placing the armature and the bag in a form tool; and conforming the
shape of the bag to the form tool. Further, this implementation may
include filling a space between the armature and the bag with air
and creating a vacuum between the form tool and the bag to force
the bag to conform to the shape of the form tool.
In accordance with another embodiment of the present invention,
applying a frame mandrel further comprises applying the frame ply
to an exterior of the bag and applying the frame mandrel over the
frame ply.
In accordance with another embodiment of the present invention,
filling further comprises compacting the media. In this embodiment,
compacting may further comprise vibrating the mandrel and frame
mandrel to aid compaction.
In accordance with another embodiment of the present invention,
applying the fiber comprises winding the fiber over the mandrel and
frame mandrel to form the structure. In this embodiment, winding
may further include placing a first winding aid on the bag; winding
the fiber over the first winding aid, the frame mandrel, and the
mandrel to form an inner skin; cutting the inner skin to remove the
first winding aids; placing a second winding aid on the inner skin;
winding the fiber over the second winding aid and inner skin to
form an outer skin; and cutting the outer skin to remove the second
winding aids. This embodiment may also include placing a core piece
on the inner skin.
In accordance with another embodiment of the present invention,
curing further comprises placing a mold around an exterior of the
structure; sealing the mold; placing the mold in a heating device;
and applying heat to the mold using the heating device. This
embodiment may also include creating a vacuum in the mandrel and
creating a vacuum in the frame mandrel.
In accordance with another embodiment of the present invention,
curing further comprises placing a mold around an exterior of the
structure; sealing the mold; placing the mold in an autoclave; and
applying pressure to the mold using the autoclave.
In accordance with an embodiment of the present invention, the
structure is a fuselage of an aircraft.
In accordance with an embodiment of the present invention, a system
for manufacturing a one-piece closed-shape structure using a
mandrel comprises: a preparing component configured to prepare the
mandrel, wherein the mandrel comprises a bag and an armature; a
first applying component configured to apply a frame mandrel to the
mandrel to form a frame for the structure; a first filling
component configured to fill the mandrel and the frame mandrel with
media; a second applying component configured to apply a curable
resign to a fiber; a third applying component configured to apply
the fiber over the mandrel and frame mandrel to form the structure;
a curing component configured to cure the structure; a removing
component configured to remove the media from the mandrel and frame
mandrel; and an extracting component configured to extract the
mandrel and frame mandrel from the structure.
In accordance with an embodiment of the present invention, a
computer-implemented method of manufacturing a one-piece
closed-shape structure using a mandrel comprises: preparing the
mandrel, wherein the mandrel comprises a bag and an armature;
applying a frame mandrel to the mandrel to form a frame for the
structure; filling the mandrel and the frame mandrel with media;
applying a curable resign to a fiber; applying the fiber over the
mandrel and frame mandrel to form the structure; curing the
structure; removing the media from the mandrel and frame mandrel;
and extracting the mandrel and frame mandrel from the
structure.
In accordance with another embodiment of the present invention, a
system for manufacturing a one-piece closed-shape structure using a
mandrel comprises: a preparing means for preparing the mandrel,
wherein the mandrel comprises a bag and an armature; an applying
means for applying a frame mandrel to the mandrel to form a frame
for the structure; a filling means for filling the mandrel and the
frame mandrel with media; a first applying means for applying a
curable resign to a fiber; a second applying means for applying the
fiber over the mandrel and frame mandrel to form the structure; a
curing means for curing the structure; a removing means for
removing the media from the mandrel and frame mandrel; and an
extracting means for extracting the mandrel and frame mandrel from
the structure.
In accordance with another embodiment of the present invention, a
one-piece closed shape structure comprises: an outer shell formed
of a composite material; and a frame formed on an interior portion
of the outer shell, the outer shell and frame being co-cured to
form the one-piece closed shape structure. In this embodiment, the
outer shell may comprise an inner and outer skin. Further, in this
embodiment, a core material may be located between the inner and
outer skin.
In accordance with another embodiment of the present invention, a
one-piece closed shape structure comprises: an outer skin formed of
a composite material; an inner skin formed of a composite material;
a frame located on an interior portion of the inner skin; and a
core material located between the inner and outer skin, wherein the
outer skin, inner skin, frame, and core material have been co-cured
to form the one-piece closed shape structure.
In accordance with another embodiment of the invention, a one-piece
airplane fuselage comprises an outer skin formed of a composite
material; an inner skin formed of a composite material; a frame
located on an interior portion of the inner skin; and a core
material located between the inner and outer skin, wherein the
outer skin, inner skin, frame, and core material have been co-cured
to form the one-piece airplane fuselage. In this embodiment, the
airplane fuselage may further comprise at least one integrally
formed flange that has been co-cured with the outer skin, inner
skin, frame, and core material. In addition, this airplane fuselage
may further comprise at least one integrally formed wing attachment
pocket that has been co-cured with the outer skin, inner skin,
frame, core material, and flange.
Additional aspects of the invention are disclosed and defined by
the appended claims. It is to be understood that both the foregoing
general description and the following detailed description are
exemplary and explanatory and are intended to provide further
explanation of the invention as claimed.
IV. BRIEF DESCRIPTION OF THE DRAWINGS
The accompanying drawings, which are incorporated in and constitute
a part of this specification, illustrate several embodiments of the
invention and, together with the following description, serve to
explain the principles of the invention.
In the drawings:
FIG. 1A is a side view of an airplane;
FIG. 1B is a partially cut away side view of an airplane
identifying certain features of the airplane, as shown in FIG.
1A;
FIG. 2A illustrates a conventional multi-piece composite
fuselage.
As shown in FIG. 2A, conventional methods used in the industry
typically construct a fuselage 200 from two or more pieces;
FIG. 2B illustrates a one-piece fuselage in accordance with an
embodiment of the present invention;
FIG. 3 is a block diagram illustrating component processes for
manufacturing a one-piece fuselage in accordance with an embodiment
of the present invention;
FIG. 4 is a perspective view of a one-piece fuselage structure
using the component processes, as shown in FIG. 3;
FIG. 5 is a block diagram illustrating the component processes for
manufacturing a one-piece integrally-stiffened fuselage in
accordance with an embodiment of the present invention;
FIG. 6 is a perspective view of a one-piece integrally-stiffened
fuselage using the components, as shown in FIG. 5;
FIG. 7 is a block diagram illustrating the components for
manufacturing a one-piece integrally stiffened fuselage by a
process in accordance with an embodiment of the invention;
FIG. 8 is a block diagram illustrating alternative embodiments for
the process of manufacturing a one-piece integrally stiffened
fuselage in accordance with the present invention;
FIG. 9A is a block diagram illustrating the processes of
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 8;
FIG. 9B is a flow diagram illustrating the internal mandrel flow
for manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A;
FIG. 9C is a flow diagram illustrating the material flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A;
FIG. 9D is a flow diagram illustrating the part flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A;
FIG. 9E is a flow diagram illustrating the mold flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A;
FIG. 9F is a flow diagram illustrating the core flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A;
FIG. 9G is a flow diagram illustrating the process of manufacturing
a one-piece fuselage in accordance with one embodiment of the
present invention, as shown in FIG. 9A-9F;
FIG. 10A illustrates tooling preparation in accordance with an
embodiment of the present invention, as shown in FIG. 9;
FIG. 10B is a cut-away view of a portion of an armature with a bag
in accordance with an embodiment of the present invention, as
described in FIG. 10A;
FIG. 11A is a perspective view of an armature and bag in a form
tool in accordance with an embodiment of the present invention, as
shown in FIG. 10A;
FIG. 11B is a cut-away view of a portion of an armature and bag in
a form tool in accordance with an embodiment of the present
invention, as shown in FIG. 11A;
FIG. 12A illustrates introducing media into a mandrel in accordance
with an embodiment of the present invention, as shown in FIG.
13;
FIG. 12B is a cut-away view of a portion of a mandrel filled with
media in accordance with an embodiment of the present invention, as
shown in FIG. 12A;
FIG. 13 is a perspective view of installing a winding shaft in a
mandrel in a form tool in accordance with another embodiment of the
present invention, as shown in FIG. 12A-12B;
FIG. 14 illustrates a close-up perspective view of a mandrel in a
form tool in accordance with an embodiment of the present
invention, as shown in FIG. 13;
FIG. 15 illustrates another perspective view of a mandrel in a form
tool in accordance with an embodiment of the present invention, as
shown in FIG. 14;
FIG. 16A is a perspective view of the mandrel prepared for lay-up
in accordance with an embodiment of the present invention, as shown
in FIGS. 12A-12B;
FIG. 16B is a cut-away view of the mandrel prepared for lay-up in
accordance with an embodiment of the present invention, as shown in
FIG. 16A;
FIG. 17 illustrates preparing an internal mandrel for filament
winding of the inner skin in accordance with another embodiment of
the present invention, as shown in FIG. 15;
FIG. 18 illustrates another perspective view of preparing the
mandrel for filament winding in accordance with an embodiment of
the present invention, as shown in FIG. 17;
FIG. 19 illustrates preparing frame mandrels to be placed on a
mandrel in accordance with an embodiment of the present invention,
as shown in FIG. 18;
FIG. 20 illustrates preparing frame materials in accordance with an
embodiment of the present invention, as shown in FIG. 9;
FIG. 21A is a perspective view of a mandrel with frame plies and
frame mandrels in place in accordance with an embodiment of the
present invention, as shown in FIG. 9;
FIG. 21B illustrates frame plies on the mandrel in accordance with
an embodiment of the present invention, as shown in FIG. 21A;
FIG. 21C illustrates frame plies and a frame mandrel on the mandrel
in accordance with an embodiment of the present invention;
FIG. 22 illustrates wing attachment plies being applied to a
mandrel to form wing attachment pockets in accordance with an
embodiment of the present invention, as shown in FIGS. 21A-21C;
FIG. 23 illustrates frame plies in frame recesses in a mandrel in
more detail in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C;
FIG. 24A illustrates a frame mandrel in a frame recess in a mandrel
in more detail in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C;
FIG. 24B illustrates a frame mandrel over frame plies in a frame
recess in a mandrel in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C, 23, and 24A;
FIG. 25 illustrates preparing the mandrel for filament winding of
the inner skin in accordance with an embodiment of the present
invention, as shown in FIG. 9;
FIG. 26 illustrates applying filament to the mandrel for filament
winding of the inner skin by a filament winding machine in
accordance with an embodiment of the present invention, as shown in
FIG. 25;
FIG. 27A is a perspective view of a mandrel with a filament-wound
inner skin in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C;
FIG. 27B is a cut-away view of a mandrel with a filament-wound
inner skin in accordance with the embodiment of the present
invention, as shown in FIG. 27A;
FIG. 28 is a side view of a mandrel with a filament-wound inner
skin with external end hoop plies in accordance with an embodiment
of the present invention, as shown in FIG. 26;
FIG. 29 illustrates cutting a mandrel in accordance with an
embodiment of the present invention, as shown in FIG. 9;
FIG. 30A is a perspective view of a mandrel with inner skin cut and
draped in accordance with an embodiment of the present invention,
as shown in FIG. 27A;
FIG. 30B is a cut-away view of a mandrel with inner skin that has
been cut and draped in accordance with an embodiment of the
invention, as shown in FIG. 30A;
FIG. 31A illustrates machining core in accordance with an
embodiment of the present invention, as shown in FIG. 9;
FIG. 31B is a perspective view of a mandrel with core material in
accordance with an embodiment of the present invention, as shown in
FIG. 9;
FIG. 31C is a cut-away view of a mandrel with core details in
accordance with an embodiment of the present invention, as shown in
FIG. 31A;
FIG. 32 illustrates a portion of a mandrel with film adhesive
covering core material in accordance with an embodiment of the
present invention, as shown in FIGS. 31A-31B;
FIG. 33 illustrates preparing a mandrel for application of an outer
skin by a filament winding machine in accordance with an embodiment
of the present invention, as shown in FIG. 9;
FIG. 34 illustrates applying an outer skin to a mandrel by a
filament winding machine in accordance with an embodiment of the
present invention, as shown in FIG. 33;
FIG. 35A is a perspective view of a mandrel with a filament wound
outer skin in accordance with an embodiment of the present
invention, as shown in FIG. 9;
FIG. 35B is a cut-away view of a mandrel with a filament wound
outer skin in accordance with an embodiment of the present
invention, as shown in FIG. 35A;
FIG. 36A is a perspective view of a mandrel with outer skin cut and
draped in accordance with an embodiment of the present invention,
as shown in FIG. 9;
FIG. 36B is a cut-away view of a mandrel with outer skin that has
been cut and draped in accordance with an embodiment of the
invention, as shown in FIG. 36A;
FIG. 37 illustrates the mandrel after cutting and draping of the
outer skin in accordance with an embodiment of the present
invention, as shown in FIG. 33;
FIG. 38A illustrates preparing a circumferential mold for a mandrel
in accordance with an embodiment of the present invention, as shown
in FIG. 9;
FIG. 38B is a cut-away view of a mandrel in the circumferential
mold in accordance with an embodiment of the present invention;
FIG. 39A illustrates preparing a circumferential mold with a vacuum
system for the frame mandrels during curing in accordance with an
embodiment of the present invention, as shown in FIGS. 38A-38B;
FIG. 39B illustrates a cut-away of the mandrel in the
circumferential mold with a vacuum system for the frame mandrels in
accordance with an embodiment of the present invention, as shown in
FIG. 39A;
FIG. 39C illustrates a vacuum port in a frame mandrel in accordance
with an embodiment of the present invention, as shown in FIGS. 39A
and 39B;
FIG. 39D illustrates a device for maintaining a vacuum in a frame
mandrel in accordance with an embodiment of the present invention,
as shown in FIGS. 39B and 39C;
FIG. 40 illustrates curing a filament wound mandrel in a
circumferential mold in an oven in accordance with an embodiment of
the present invention, as shown in FIG. 9;
FIG. 41 illustrates removing a circumferential mold from around a
one-piece integrally stiffened fuselage on a mandrel in accordance
with an embodiment of the present invention, as shown in FIG.
9;
FIG. 42 illustrates removing media from a mandrel in accordance
with an embodiment of the present invention, as shown in FIG.
9;
FIG. 43 illustrates a one-piece integrally stiffened fuselage
contained in a circumferential mold after removal of media and
armature in accordance with one embodiment of the present invention
as shown in FIG. 42;
FIG. 44 illustrates removing a bag from a one-piece integrally
stiffened fuselage in accordance with an embodiment of the present
invention as shown in FIG. 41;
FIG. 45 illustrates a bag after removal from a mandrel in
accordance with an embodiment of the present invention, as shown in
FIG. 44;
FIG. 46 illustrates removing frame mandrels from a one-piece
integrally stiffened fuselage in accordance with an embodiment of
the present invention as shown in FIG. 9;
FIG. 47 illustrates a one-piece integrally-stiffened fuselage
manufactured in accordance with one embodiment of the present
invention as shown in FIG. 9;
FIG. 48 illustrates a one-piece integrally stiffened fuselage
manufactured in accordance with one embodiment of the present
invention as shown in FIG. 9;
FIG. 49 is a block diagram illustrating the process of
manufacturing a one-piece fuselage in accordance with another
embodiment of the present invention, as shown in FIG. 9;
FIG. 50 illustrates assembling a circumferential mold around a
mandrel in accordance with an embodiment of the present invention,
as shown in FIG. 49;
FIG. 51 illustrates bagging a circumferential mold in accordance
with an embodiment of the present invention, as shown in FIG.
50;
FIG. 52 illustrates placing a circumferential mold in an autoclave
for curing in accordance with an embodiment of the present
invention, as shown in FIG. 51;
FIG. 53 illustrates removing a circumferential mold after curing in
an autoclave in accordance with an embodiment of the present
invention, as shown in FIG. 52; and
FIG. 54 illustrates a one-piece integrally-stiffened fuselage
manufactured in accordance with another embodiment of the present
invention, as shown in FIG. 49.
V. DESCRIPTION OF THE EMBODIMENTS
A. Introduction
Methods and structures in accordance with the present invention
will now be described with respect to an embodiment of a one-piece
structure, an aircraft fuselage. The attached figures illustrate
the manufacture of both a fuselage containing a tail cone and a
fuselage without a tail cone. The invention as claimed, however, is
broader than fuselages and extends to other closed-shape
structures, such as, other aircraft, automotive, forklift, or
watercraft structures.
B. Methods and Structures
FIG. 1A is a side view of an airplane. As shown in FIG. 1A,
airplane 100 consists of engine section 101, fuselage 102,
empennage 103, and wings 104. Engine section 101, empennage 103,
and wings 104 connect to fuselage 102. Airplane 100 may be any type
of airplane, such as, prop, jet, or other type. Airplane 100 is the
type of airplane for which a one-piece fuselage could be
constructed (which is described in more detail below). In one
implementation, the one-piece fuselage includes fuselage 102. In
other implementations, the one-piece fuselage may also include
engine section 101, empennage 103, wings 104, and/or any other
parts of aircraft 100 (not shown). This implementation is merely
exemplary, and other implementations may also be used.
FIG. 1B is a partially cut away side view of an airplane
identifying certain features of the airplane, as shown in FIG. 1A.
In FIG. 1B, airplane 100 is described in more detail than in FIG.
1A. As shown in FIG. 1B, engine section 101 contains an engine 107,
an engine mount 108, and a firewall 109. Engine section 101, engine
mount 108, and firewall 109 are all connected to fuselage 102. In
some implementations, engine 107 is connected to fuselage 102 via
engine mount 108. However, other implementations may have engine
107 connected directly to fuselage 102. Further, engine mount 108
may be a separate component, as shown in FIG. 1B, or engine mount
108 may be a part of either engine section 101 or fuselage 102.
These implementations are merely exemplary, and other
implementations may also be used.
Empennage 103 contains tail cone 106, vertical stabilizers 107, and
horizontal stabilizers 108. Empennage 103 may be a separate
component of airplane 100, as shown in FIG. 1B, or tail cone 106
may be a part of fuselage 102 with vertical stabilizer 107 and
horizontal stabilizer 108 being separate pieces. These
implementations are merely exemplary, and other implementations may
also be used.
Wings 104 generally include left wing 121 (shown) and right wing
122 (not shown). Wings 104 are connected to fuselage 102 by wing
spars 110. Wing spars 110 support wings 101 within fuselage 102.
Other wing configurations may be used for airplane 100, such as a
bi-wing configuration or a tri-wing configuration or other wing
configurations. In addition, a canard (not shown) and winglets (not
shown) may also be used with airplane 100. Also, airplane 100
depicts a low wing aircraft, but airplane 100 may also be a
high-wing, mid-wing, or other wing-design aircraft.
Fuselage 102 contains panel section 116, seat supports (not shown),
seats 112, access doors 115, luggage access doors 113, and windows
114. Panel section 116 holds flight instruments for airplane 100.
Seat supports hold seats 112. Access door 115 is depicted as a
single door on the left side of airplane 100, as shown in FIG. 1B.
Access door 115 may also be located on the right side of airplane
100. Further, additional or other doors may be included within a
group of access doors 115, such as a second set of access doors or
any other access door configurations. Luggage access door 113 is
depicted as located on the back left side of airplane 100, as shown
in FIG. 1B, but luggage access door 113 may be located anywhere on
airplane 100. In addition, airplane 100 may contain multiple
luggage access doors 113. Fuselage 102 also contains wing spar
attachment boxes (not shown).
Windows 114 include front window 117 and side windows 118. FIG. 1B
depicts two side windows 118, but other configurations may be used
for side windows 118 such as one side window 118 or two or more
side windows 118. Windows 114 may also include a rear window (not
shown). Windows 114 may also include other windows, such as
skylight windows (not shown), camera windows (not shown), or any
other type of window.
As shown in FIG. 1B, fuselage 102 has numerous openings, such as
access doors 115, windows 114, and luggage access doors 113. Also,
there are other openings that are not shown such as an engine mount
block (not shown) for engine mount 108, an empennage mounting block
(not shown) for empennage 106, and landing gears mounts (not shown)
for landing gear 119.
Because of these openings, portions of fuselage 102 may be
strengthened for support around these openings. For example, window
frames 140 may be strengthened to support windows 114. Other areas
may also be strengthened, such as seat supports 130 (not shown, but
described above). Other strengthening may also be necessary for the
engine mount block (not shown), the landing gear mounts (not
shown), the empennage mounting block (not shown), and roll-over
frames (not shown). Still other areas may also need to be
strengthened, depending on the design of airplane 100. These
implementations are merely exemplary, and other implementations may
also be used.
FIG. 2A illustrates a conventional multi-piece composite fuselage.
As shown in FIG. 2A, conventional methods used in the industry
typically construct a fuselage 200 from two or more pieces. Those
pieces consist of fuselage halves 205 and 210, frame stiffening
structures for the passenger area 220, and wing spar attachment
boxes 230. In composite aircraft manufacture, these pieces are
typically manufactured from fiberglass prepreg. These conventional
methods also require the steps of bonding the pieces together,
machining of the pieces at the joint areas, machining the core
frames, and various other mechanical assembly processes.
FIG. 2B illustrates a one-piece fuselage in accordance with an
embodiment of the present invention. As shown in FIG. 2B, a
fuselage 250 is a one-piece structure, including exterior surface
of the fuselage 260, frame sections 270, attachment pockets for the
wings 280, and other frames sections, attachments pockets, and
flanges (not shown). Notably, the use of a one-piece fuselage
eliminates the assembly operations that are associated with the
conventional methods for manufacturing a fuselage as well as
providing other advantages, as described above. This implementation
is merely exemplary, and other implementations may also be
used.
FIG. 3 is a block diagram illustrating component processes for
manufacturing a one-piece fuselage in accordance with an embodiment
of the present invention. As shown in FIG. 3, the component
processes for manufacturing a one-piece fuselage 300 include
molding 310, tooling 320, tooling integration 330, and fiber
placement 340. Molding 310 includes the use of any type of molding.
For example, molding 310 may include such things as hand lay up of
fiberglass or graphite prepreg into molds, pressing of sheet
molding compounds, injection of molding compounds into dies, and/or
machine lamination of composite prepreg onto molds.
Tooling 320 includes the use of any type of tooling needed for
molding. For example, tooling 320 may include the use of metal
molds, molds made from composite materials, and/or mandrels made
from metals and composite materials. Tooling 320 also includes
tooling made from elastomeric materials such as silicone, urethane,
or natural rubbers. Tooling 320 further includes the use of plastic
or metal dies and punches.
Tooling integration 330 includes any combination of molding 310
with tooling 320. For example, tooling integration 330 includes
vacuum sealing of a part cavity, pressurization of tool cavities,
and/or application of vacuum pressure in tool cavities.
Fiber placement 340 includes any placement using any form of fiber.
For example, fiber placement 340 includes such things as winding
with carbon tape, winding with carbon tow, winding with glass fiber
or roving, winding with glass tape, wrapping of glass or carbon
prepreg materials, and/or wrapping of carbon and glass fiber
materials. As shown in FIG. 3, molding 310, tooling 320, tooling
integration 330, and fiber placement 340 may be used to create
one-piece fuselage 300. This implementation is merely exemplary,
and other implementations may also be used.
FIG. 4 is a perspective view of a one-piece fuselage structure
using the component processes, as shown in FIG. 3. As shown in FIG.
4, a one-piece fuselage 400 may be created using the processes
described in FIG. 3. For example, molding 310 is used to create
such things as molded frames 410, molded integral flanges 415 (for
attachment of the bulkhead), and molded integral wing attachment
hard points 420 (for attachment of the wings). Tooling 320 is used
to create such things as armature 425 and mandrel 430 (placed
inside the fuselage). Tooling integration 330 is used to integrate
tooling 320 with molding 310 to prepare for fiber placement 340 of
fuselage 400. Fiber placement 340 creates fuselage skin 450. FIG. 4
depicts just some examples of the uses of the components of FIG. 3
in a one-piece fuselage, and many other uses may be made of these
components just some of which are described herein with reference
to FIG. 4).
FIG. 5 is a block diagram illustrating the component processes for
manufacturing a one-piece integrally-stiffened fuselage in
accordance with an embodiment of the present invention. As shown in
FIG. 5, a one-piece integrally-stiffened fuselage 500 includes
molding co-cured hollow and foam-filled frames 510, elastomeric
tooling for internal fuselage mandrel 520, tooling integration for
integrally-stiffened fuselage 530, and fiber placement fuselage
shape 540.
For fuselage 500, molding co-cured hollow and foam filled frames
510 includes the molding of stiffening structure inside of the
fuselage shell that is co-cured with that shell. Such molded
structure may also include flanges that are integral with the
shell. Other molded structures may further include wing attachment
pockets and pockets for engine truss mount fittings.
Elastomeric tooling for internal fuselage mandrel 520 includes the
use of an elastomeric tooling associated with molding the internal
shape of the fuselage. In this context, elastomeric tooling refers
to a mandrel that is used to maintain the internal shape of the
fuselage during frame lay up and filament winding.
Tooling integration for integrally-stiffened fuselage 530 involves
the joining of molding 510 and elastomeric tooling 520 in a manner
that produces the fuselage shape. Tooling integration 530 includes
such things as application of a vacuum and/or pressure in various
mold cavities to obtain the desired fuselage shape.
Finally, fiber placement fuselage shape 540 is used to create the
fuselage skin material and shape. During fiber placement 540, fiber
is wound directly on the elastomeric mandrel to create the fuselage
skin. Fiber placement 540 also includes the creation of skin
material by winding fiber over a secondary mandrel and then cutting
ply pieces to obtain frame and integral stiffening structures.
Co-curing of fuselage components creates integral stiffening for
one-piece integrally stiffened fuselage 500. The co-curing of these
components reduces the chances of defects created during
manufacturing of separate elements and their subsequent joining
using bonding or mechanical fasteners. The co-cured integral
stiffening distributes loads more uniformly throughout the fuselage
during flight and in the event of an off field landing.
FIG. 6 is a perspective view of a one-piece integrally-stiffened
fuselage using the components, as shown in FIG. 5. As shown in FIG.
6, a one-piece fuselage 600 may be created using the components
described in FIG. 5. For example molding 510 may be used to create
co-cured hollow and foam filled frames 610, integral flanges 615
(such as for bulkhead attachment), and integral wing attachment
hard points 620.
Tooling 520 may be used to create armature 625 and a reusable
elastomeric mandrel 630 both of which go inside of fuselage 600.
Tooling 520 may also include filament-wound broad goods. In
general, broad goods are custom-sized pieces of composite
materials, and filament-wound broad goods are custom-sized pieces
of composite materials that have been filament-wound. For example,
a filament-wound broad good would become a co-cured hollow frame
ply. These filament-wound broad goods may be used for integral
frames 610, flanges 615, longerons 610, and doublers 620.
Tooling integration 530 may be used to integrate such things as
tooling for fuselage 600 and tooling for the filament winding of
fuselage skins (such as fuselage skin 650) with molding 510.
Fiber placement 540 may be used to create fuselage skin plies 650.
FIG. 6 depicts just some examples of the uses of the components of
FIG. 5 in a one-piece integrally-stiffened fuselage, and many other
uses may be made of these components Oust some of which are
described with reference to FIG. 6).
FIG. 7 is a block diagram illustrating the components for
manufacturing a one-piece integrally stiffened fuselage by a
process in accordance with an embodiment of the invention. As shown
in FIG. 7, process for creation of one-piece integrally-stiffened
fuselage 700 includes molding co-cured hollow and foam filled
frames 510, elastomeric tooling for internal fuselage mandrel 520,
tooling integration 530, and fiber placement fuselage shape 540. As
shown in FIG. 7, a process for the combination of molding co-cured
hollow and foam filled frames 510, elastomeric tooling for internal
fuselage mandrel 520, tooling integration 530, and filament
placement fuselage shape 540 may result in one-piece
integrally-stiffened fuselage 500. This implementation is merely
exemplary, and other implementations may also be used.
FIG. 8 is a block diagram illustrating alternative embodiments for
the process of manufacturing a one-piece integrally stiffened
fuselage in accordance with the present invention. As shown in FIG.
8, process for creation of one-piece integrally-stiffened fuselage
(as described in FIG. 7) includes Process Alternate #1 805, Process
Alternate #2 810, and Other Process Alternates 815. Other Process
Alternates 815 shows that various alternative processes may be used
for creating one-piece integrally-stiffened fuselage 700. Process
Alternate #1 805 is depicted in FIGS. 9-48. Process Alternate #2
810 is depicted in FIGS. 49-53. Although not depicted in separate
figures, Other Process Alternates 815 show that processes other
than those described herein may be used for the process for
creation of one-piece integrally-stiffened fuselage 800.
1. Alternate 1
FIG. 9A is a block diagram illustrating the processes of
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 8. As shown
in FIG. 9A, process for creation of one-piece integrally stiffened
fuselage 900 includes five processes including internal mandrel
flow 901, material flow 902, part flow 903, mold flow 904, and core
flow 905. Internal mandrel flow 901 is depicted in FIG. 9B.
Material flow 902 is depicted in FIG. 9C. Part flow 903 is depicted
in FIG. 9D. Mold flow 904 is depicted in FIG. 9E. Core flow 905 is
depicted in FIG. 9F. FIG. 9G illustrates the integration of flows
901-905.
FIG. 9B is a flow diagram illustrating the internal mandrel flow
for manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A. As shown
in FIG. 9B, internal mandrel flow 901 describes the flow of the
internal tooling for creation of the fuselage. Internal mandrel
flow 901 begins with tooling preparation 907. Tooling preparation
907 includes placing of an armature inside of a bag, forming the
bag inside of a form tool, and filling the space between the bag
and the form tool with media to form the internal mandrel. The
mandrel exterior (which is described in detail on the following
figures) is in the shape of the desired fuselage interior.
After tooling preparation 907, the following actions take place:
place frames and frame mandrels 908, wind inner skin 910, cut and
drape 911, place core 912, wind outer skin 913, cut and drape 914,
close mold 915, cure 916, demold 917 and extract frame mandrels
918. Each of these actions are described in detail below (see FIGS.
10-48).
FIG. 9C is a flow diagram illustrating the material flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A. As shown
in FIG. 9C, material flow 902 describes the flow of fiber and resin
for creating the composite material that becomes one-piece
integrally-stiffened fuselage 500. Material flow 902 begins with
inspect incoming materials 930. Inspect incoming materials 930
involves inspection of the fiber and resin used by system 900.
Fiber and resin at inspect incoming materials 930 are inspected for
conformity for use in prepare frame materials 932 and filament
winder 935.
Prepare frame materials 932 involves the preparation of a fiber and
resin, which includes filament winding of broad goods, cutting with
a ply cutter, and the preparation of frame and doubler plies.
Filament winder 935 includes loading of the fiber and resin into a
filament winder device, such as those known in the art. Filament
winder 935 creates an inner skin in wind inner skin 910. Filament
winder 935 also creates an outer skin in wind outer skin 913. Each
of these actions is described in detail below (see FIGS.
10-48).
In material flow 902, any type of fiber and any type of resin may
be used. Some of the fibers that are found to be acceptable
include: Toray T700, T600, and T300 that are available in 3 K, 6 K,
and 12 K tow count, Amoco T-300 and T-650 that are available in 3
K, 6 K, and 12 K tow count, Hexcal AS4 that is available in 3 K and
6 K tow count, Fortafil 3(C), Grafil 34-600WD, and Panex 33. In one
implementation, fibers that are never twisted may be used, although
other fibers may also be used in other implementations. In most
implementations, an acceptable tow count for the spool is dependent
upon part size.
Additionally, any type of curable resin may be used. Some curable
resins that have been found acceptable include epoxy resin with a
room temperature viscosity of 10,000 to 125,000 cps. The viscosity
to be used generally depends upon the shape of the part being
filament wound. Any type of Shell epoxy resin may also be used.
Shell epoxy resins that have been found acceptable include
combinations of 862 and 1050 with "W" curative and accelerator 537.
In addition, Shell epoxy resins that have been found acceptable may
use tougheners from Nippon zeon, which have been shown to have
desirable physical properties. Shell epoxy resins can be used
separately or in combination with other resins to obtain the
desired properties. Furthermore, still other resins may also be
used. For example, high tack resins may be used under certain
circumstances (such as holding fibers in place during winding).
Therefore, any type of curable resin and any combination of curable
resin types may be used.
FIG. 9D is a flow diagram illustrating the part flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A. As shown
in FIG. 9D, part flow 903 describes the flow of the part (i.e. the
fuselage) during the manufacturing process. Part flow 903 includes
most of the actions of internal mandrel flow 901. Part flow 903
begins with prepare frame materials 932. Part flow 903 then
includes the following actions: place frames and frame mandrels
908, wind inner skin 910, cut and drape 911, place core 912, wind
outer skin 913, cut and drape 914, close mold 915, cure 916,
de-mold 917, extract frame mandrels 918, visually inspect part 919,
trim 920, prime and paint 921, and store for assembly 922. Part
flow 903 also includes monitoring process 936, monitoring process
937, and monitoring process 978 for monitoring wind inner skin 910,
wind outer skin 913, and cure 916, respectively. Finally, part flow
903 also includes prepare oven 975, and cut-outs to quality control
985. Each of these actions is described below (see FIGS.
10-48).
FIG. 9E is a flow diagram illustrating the mold flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A. As shown
in FIG. 9E, mold flow 904 describes the actions involving a mold
during the manufacturing process. In one implementation, a mold
(which will be described in more detail below) is in the shape of a
circumferential mold for holding (therefore molding) the structure
inside the circumferential mold. In one implementation, the
circumferential mold is made of metal. In other implementations,
the circumferential mold may be made out of any other materials.
Mold flow 904 begins with prepare mold 940. After prepare mold 940,
close mold 915 occurs. Close mold 915 includes closing the mold. In
some implementations, close mold 915 may also include applying a
vacuum to the mold or applying pressure to the mold. After close
mold 915, cure 916 occurs (which in some implementations may be
preceded by prepare oven 975). In one implementation, cure 916
involves heating the mold, or in another implementation cure 916
involves pressurization of the mold. In other implementations, cure
916 may include curing by any other method. After cure 916, demold
917 occurs. Demold 917 includes the removal of the contents of the
mold. After demold 917, the mold is reused in prepare mold 940. All
of these actions will be described below (see FIGS. 10-48).
FIG. 9F is a flow diagram illustrating the core flow for
manufacturing a one-piece fuselage in accordance with one
embodiment of the present invention, as shown in FIG. 9A. As shown
in FIG. 9F, core flow 905 involves the core materials that give
buckling stiffness to the desired structure, such as one-piece
integrally-stiffened fuselage 500. Core flow 905 begins with
inspect incoming materials 930. At inspect incoming materials 930,
core material (such as honeycomb core material) is inspected for
use. After inspect incoming materials 930, the core material is
machined to shape in machine core to shape 952. After machine core
to shape 952, the core material is cleaned in clean core 950. After
clean core 950, the core material is used by system 900 (See FIG.
9G) in place core 912. Notably, place core 912 may involve
manipulation of the shape of the desired structure by cut and drape
911. For example, between cut and drape 911 and place core 912,
additional material may be added to the core material for
manipulation of the shape of the desired structure. Any type of
core material may be used for the core material. These actions will
be described below (see FIGS. 10-48).
FIG. 9G is a flow diagram illustrating the process of manufacturing
a one-piece fuselage in accordance with one embodiment of the
present invention, as shown in FIG. 9A-9F. FIG. 9G illustrates the
combination in system 900 of internal mandrel flow 901, material
flow 902, part flow 903, mold flow 904, and core flow 905,
described in FIGS. 9A-9G. Internal mandrel flow 901 is depicted in
FIG. 9G as a line with arrows. Material flow 902 is depicted in
FIG. 9G as a line with arrows with circles through the line. Part
flow 903 is depicted in FIG. 9G as a bold line with arrows. Mold
flow 904 is depicted in FIG. 9G as a line with arrows with circles
next to the line. Core flow 905 is depicted in FIG. 9G as a line
with arrows with slash marks through the line.
As referenced above, the actions taken in system 900 are described
in FIGS. 10-48. FIGS. 10-48 illustrate the use of system 900 to
create both a one-piece integrally stiffened fuselage with a tail
cone and an integrally stiffened fuselage without a tail cone.
Therefore, FIGS. 10-48 illustrate the components of system 900
being used to create multiple structures. Those figures
illustrating the creation of a one-piece integrally stiffened
fuselage with a tail cone will be numbered consistently with one
another and those figures illustrating the creation of a one-piece
integrally stiffened fuselage without a tail cone will be number
consistently with one another.
As shown in FIG. 9G, system 900 begins with tooling preparation
907. Tooling preparation 907 is described in FIG. 10A.
FIG. 10A illustrates tooling preparation in accordance with an
embodiment of the present invention, as shown in FIG. 9. FIG. 10A
provides an example of tooling preparation 907 from FIG. 9. FIG.
10A shows the preparation of tooling, such as a mandrel 1000. In
one implementation, mandrel 1000 is a reusable elastomeric mandrel,
such as that currently available through International Design
Technologies, Inc (IDT). However, any mandrel may be used.
Mandrel 1000 may include a bag 1010 and an armature 1020. Bag 1010
may comprise premolded silicone, or bag 1010 may consist of any
other form or substance. Some silicone bag materials that have been
found acceptable include those available from Arlon-Mosite,
Kirkhill, and D-Aircraft Products SMC 950. In addition, there are
many other suppliers of high temperature (up to 400.degree. F.),
unfilled, and uncured silicone sheet materials that may be used,
depending upon the cure temperature of the desired part.
Armature 1020 may be made of any material. In one implementation, a
welded metal armature is used. However, other materials could be
used to form the armature. To minimize weight and mandrel bending,
armature 1020 may be as large as possible, while allowing it to be
removed from bag 1010 and from the completed fuselage. These
implementations are merely exemplary, and other implementations may
also be used.
FIG. 10B is a cut-away view of a portion of an armature with a bag
in accordance with an embodiment of the present invention, as
described in FIG. 10A. As shown in FIG. 10B, armature 1020 is
placed through bag 1010 to form cavity 1030. The space difference
between armature 1020 and bag 1010 provides for cavity 1030. To
form cavity 1030, bag 1010 is sealed at each end to armature 1020.
In one implementation, clamps and/or bolts are used to seal each
end of bag 1010. Armature 1020 thus supports bag 1010. Notably, bag
1010 has a desired pre-molded shape 1040. Bag 1010 may lack the
rigidity to maintain desired shape 1040 without support from
additional tooling. Therefore, as described below, additional
tooling may be used to maintain desired shape 1040. These
implementations are merely exemplary, and other implementations may
also be used.
FIG. 11A is a perspective view of an armature and bag in a form
tool in accordance with an embodiment of the present invention, as
shown in FIG. 10A. As shown in FIG. 11A, following placement of
armature 1020 in bag 1010 (as described in FIG. 10A), armature 1020
and bag 1010 are placed in a form tool 1110 and bag 1010 is sealed
at both ends to form tool 1110. In one implementation, form tool
1110 covers most of armature 1020 and bag 1010. Form tool 1110
provides a desired shape to outside surface of bag 1010. This
implementation is merely exemplary, and other implementations may
also be used.
FIG. 11B is a cut-away view of a portion of an armature and bag in
a form tool in accordance with an embodiment of the present
invention, as shown in FIG. 1A. As shown in FIG. 11B, bag 1010 is
between form tool 1110 and armature 1020. Bag 1010 is sealed at
each end to both form tool 1110 and armature 1020 to form enclosed
cavity 1030 and enclosed cavity 1120. Enclosed cavity 1030 is
between outside surface of armature 1020 and inside surface of bag
1010 and enclosed cavity 1120 is between inside surface of form
tool 1110 and outside surface of bag 1010. In one implementation,
form tool 1110 is equipped with ports for enclosed cavity 1120 (not
shown) to control pressure and quantity of air within enclosed
cavity 1120. Enclosed cavity 1030 may also be equipped with ports
(not shown) to control pressure and quantity of air within enclosed
cavity 1030. These implementations are merely exemplary, and other
implementations may also be used.
In one implementation, to provide the desired shape to outside
surface of bag 1010, the air is vented from enclosed cavity 1120
through ports 1130 (not shown) while pressurized air is inserted
into enclosed cavity 1030 through ports (not shown) forcing outside
surface of bag 1010 against inside surface of form tool 1110. Ports
for enclosed cavity 1120 are then sealed to maintain outside
surface of bag 1010 against inside surface of form tool 1110. Ports
to enclosed cavity 1030 may then be kept pressurized or they may be
vented to the atmosphere. In another implementation, to provide
desired shape to outside surface of bag 1010, the air is evacuated
from enclosed cavity 1120 through ports while ports into enclosed
cavity 1030 are left vented to the atmosphere which forces outside
surface of bag 1010 against inside surface of form tool 1110. These
implementations are merely exemplary, and other implementations may
also be used.
FIG. 12A illustrates introducing media into a mandrel in accordance
with an embodiment of the present invention, as shown in FIG. 13.
As shown in FIG. 12A, after forming bag 1010 to the shape of form
tool 1110 (as described in FIGS. 11A-11B), media 1210 may be
introduced into enclosed cavity 1030 to provide the desired shape
to bag 1010 (not shown, but shown in FIG. 12B). In one
implementation, media 1210 is introduced through a sealable opening
(not shown) inside armature 1020. Media 1210 may be any material
used to provide rigidity to bag 1010. In one implementation, media
1210 is a lightweight insulator material, such as porous ceramic
materials used for water filtration. In another implementation,
aluminum hollow-beaded materials may be used. These implementations
are merely exemplary, and other implementations may also be
used.
In one implementation, when media 1210 is introduced into enclosed
cavity 1030, it may be introduced under pressure if enclosed cavity
1030 is pressurized, under atmospheric conditions if enclosed
cavity 1030 is vented to atmosphere, or under less than atmospheric
conditions if enclosed cavity 1030 is maintained under some
pressure less than atmospheric. As shown in FIG. 12A, in one
implementation, the introduction of media 1210 is performed in a
semi-horizontal orientation. However, in other implementations,
other orientations, such as a vertical orientation or any other
orientation, may be used for introducing media 1210 into enclosed
cavity 1030. These implementations are merely exemplary, and other
implementations may also be used.
FIG. 12B is a cut-away view of a portion of a mandrel filled with
media in accordance with an embodiment of the present invention, as
shown in FIG. 12A. As shown in FIG. 12B, media 1210 is introduced
into enclosed cavity 1030, which is between armature 1020 and bag
1010, as held together by form tool 1110. After the introduction of
media 1210, media 1210 may be compacted to settle the media. In one
implementation, the compacting of media 1210 occurs by vibrating
form tool 1110. In another implementation, compacting of media 1210
occurs by tamping media 1210. These implementations are merely
exemplary, and other implementations may also be used.
Following compaction of media 1210 the air within enclosed cavity
1030 may be removed as completely as possible to complete a
pressure difference between enclosed cavity 1030 and the
atmosphere. This pressure difference causes bag 1010 to retain its
shape once form tool 1110 is removed. If a pressure difference
between enclosed cavity 1030 and the atmosphere is not maintained,
bag 1010 may lose the desired shape established by form tool 1110.
In one implementation, five pounds per square inch (psi) of
pressure difference between enclosed cavity 1030 and atmospheric
pressure has been demonstrated sufficient to cause bag 1010 to
retain the desired shape. This implementation is merely exemplary,
and other implementations may also be used.
FIG. 13 is a perspective view of installing a winding shaft in a
mandrel in a form tool in accordance with another embodiment of the
present invention, as shown in FIGS. 12A-12B. FIG. 13 is also the
first drawing illustrating the manufacture of a mandrel without a
tail cone. As shown in FIG. 13, following the introduction of media
into mandrel 1390, a winding shaft 1330 is inserted into mandrel
1300. In one implementation, mandrel 1390 incorporates armature
1320, winding shaft 1330, compacted media (not shown), and bag (not
shown). Winding shaft 1330 is used to rotate mandrel 1390 during
fiber placement. In one implementation, winding shaft 1330 is
inserted into a box channel within armature 1320.
As shown in FIG. 13, mandrel 1390 is surrounded by form tool 1310,
as described above. Form tool 1310 incorporates pivot 1340. In one
implementation, pivot 1340 allows form tool 1310 and mandrel 1300
to rotate to vertical, if needed. Form tool 1310 also includes
clamps 1350, bolts 1360, vacuum port 1370, and end plates 1380. In
one implementation, clamps 1350 are used to seal a bag (not shown)
around form tool 1310. In this implementation, bolts 1360 are used
to join and seal segments of form tool 1310 to each other. In
addition, bolts 1360 may also be used to seal end plates 1380 to
bag 1420, form tool 1310 and to armature 1320.
FIG. 14 illustrates a close-up perspective view of a mandrel in a
form tool in accordance with an embodiment of the present
invention, as shown in FIG. 13. As shown in FIG. 14, one section of
form tool 1310 has been removed from around mandrel 1390. As shown
in FIG. 14, the external surface of bag 1420 is formed to the shape
of internal surface of form tool 1310.
In one implementation, form tool 1310 includes a vacuum port 1430.
Vacuum port 1430 connects to an interior surface of form tool 1310.
Vacuum port 1430 is used to vent or remove air from between
interior surface of form tool 1310 and exterior surface of bag
1420.
FIG. 15 illustrates another perspective view of a mandrel in a form
tool in accordance with an embodiment of the present invention, as
shown in FIG. 14. As shown in FIG. 15, a section of form tool 1310
has been removed from mandrel 1390. As further shown in FIG. 15,
bag 1420 retains the desired shape imparted to it by form tool
1310. In one implementation, a pressure differential is maintained
between a media-filled enclosed cavity situated between armature
1320 and bag 1420 and the atmosphere. This implementation is merely
exemplary, and other implementations may also be used.
As shown in FIG. 15, mandrel 1390 includes frame recesses 1540 and
wing attachment pocket recesses 1550. Frame recesses 1540 and wing
attachment pocket recesses 1550 are located on the external surface
of bag 1420. In one implementation, frame recesses 1540 and wing
attachment pocket recesses 1550 are created by the inside surface
of form tool 1310. However, it may be problematic for form tool
1310 to create frame recesses 1540 and wing attachment pocket
recesses 1550 because of the tendency of form tool 1310 to have
either no draft or negative draft. For this reason, the removal of
form tool 1310 could be difficult from around certain portions of
mandrel 1390.
In one implementation, these problems are overcome by making form
tool 1310, as shown in FIG. 15, in multiple pieces having required
draft. In another implementation, the negative draft features are
made as separate details that fit within recesses in the inside of
the form tool and are detachable from outside of the form tool when
it is necessary to remove form tool from around formed mandrel.
These implementations are merely exemplary, and other
implementations may also be used.
FIG. 16A is a perspective view of the mandrel prepared for lay-up
in accordance with an embodiment of the present invention, as shown
in FIGS. 12A-12B. As shown in FIG. 16A, following introduction of
internal media 1210 (described in FIGS. 12A-12B), form tool 1110 is
removed to expose mandrel 1610. Mandrel 1610 is then cleaned and
prepared for lay-up. Lay-up is the procedure of applying composite
materials at desired locations to the exterior surface of the
formed mandrel. These materials may (when cured) form stiffening
structure, frames, within the fuselage, or when placed following
placement of an inner skin, as core details, add buckling strength
to the fuselage skin.
In one implementation, mandrel 1610 contains frame recesses 1620,
window recesses 1630, door recesses 1640, core detail recesses
1650, and wing attachment pocket recesses 1660. These recesses are
used to form features such as frames, windows, doors, core pockets,
and wing attachment pockets in the fuselage. This implementation is
merely exemplary, and other recesses and other implementations may
also be used.
FIG. 16B is a cut-away view of the mandrel prepared for lay-up in
accordance with an embodiment of the present invention, as shown in
FIG. 16A. As shown in FIG. 16B, mandrel 1610 includes armature
1020, bag 1010, and enclosed cavity 1030 filled with media 1210. As
shown in FIG. 16B, the insertion of media 1210 into enclosed cavity
1030 and subsequent evacuation of air has caused the outside
surface of bag 1010 to hold the desired inside surface shape of the
removed form tool 1110. This "shape memory" provides for features
desired for lay-up of window doublers, door frames, core details,
and wing attachment lugs. This implementation is merely exemplary,
and other implementations may also be used.
FIG. 17 illustrates preparing an internal mandrel for filament
winding of the inner skin in accordance with another embodiment of
the present invention, as shown in FIG. 15. As shown in FIG. 17,
mandrel 1390 is placed in a winding cart 1710 in preparation for
filament winding. Winding end aids 1720 are positioned on the ends
of mandrel 1500. Winding end aids 1720 are used to eliminate
concave or flat winding surfaces. Other winding aids are also
depicted in FIG. 17. Examples of other winding aids include frames
1740, door recess fillers 1760, and windshield area fillers 1750.
Other winding aids provide a surface upon which the filament
winding machine places the fibers (and resin) so that the fibers do
not shift as the mandrel is rotated. Other winding aids 1725 are
also used to protect mandrel 1390 from being cut during cut and
drape 911. Other winding aids 1725 may further include guide
features to guide cutting of plies during cut and drape 911
(described below). In addition, gap winding aids 1727 (not shown)
are also used in areas where other winding aids 1725 are higher
than mandrel 1390. These gap winding aids 1727 ensure that all
surfaces are convex prior to filament winding.
FIG. 18 illustrates another perspective view of preparing the
mandrel for filament winding in accordance with an embodiment of
the present invention, as shown in FIG. 17. As shown in FIG. 18,
gap winding aids 1727 have been installed to make all surfaces
convex. Gap winding aids 1727 are used because the excessive
concave area on mandrel 1390 would make filament winding
difficult.
FIG. 19 illustrates preparing frame mandrels to be placed on a
mandrel in accordance with an embodiment of the present invention,
as shown in FIG. 18. As shown in FIG. 19, frame mandrel tools 1910
are used to create desired shapes using frame mandrels 1920. In one
implementation, frame mandrels 1920 are held to the inside shape
frame mandrel tools 1910 by drawing a vacuum between the inside
surface of form tools 1910 and the outside surface of frame
mandrels 1920. Frame mandrels 2020 are then filled with media.
Finally, this media is compacted to give the frame mandrel the
desired shape.
In one implementation, to maintain the desired shape, the air
within the frame mandrel cavity, which has been completely filled
with media, is evacuated. This causes media to lock together
retaining the form tool shape. In another implementation, frame
mandrels may include armatures. These implementations are merely
exemplary, and other implementations may also be used.
FIGS. 10-19 have described tooling preparation 907, as shown in
FIG. 9. As shown in FIG. 9, following tooling preparation 907,
prepare frame materials 932 occurs. Prepare frame materials 932 is
described in FIG. 20.
FIG. 20 illustrates preparing frame materials in accordance with an
embodiment of the present invention, as shown in FIG. 9. As shown
in FIG. 20, frame material 2000 may be cut to produce ply pieces
which will be formed and placed on a mandrel in a predetermined
location to produce frames. Frame material 2000 may also be cut to
produce ply pieces, which will be formed and placed on a mandrel in
predetermined locations to produce integral doublers, longerons,
flanges, and attachment lugs.
Frame material 2000 may be used for frame plies 2010, doubler plies
2020, longeron plies 2030, integral flange plies 2040, and wing
attachment pocket plies 2050. Frames, doublers, longerons, flanges,
and attachment pockets are structures that enhance the strength and
utility of the fuselage. Frame material 2000 may include prepreg
fabric or filament-wound broad goods. Other frame material may also
be used.
FIG. 20 has described prepare frame materials 932, as shown in FIG.
9. As shown in FIG. 9, following prepare frame materials 932, place
frames and frame mandrels 908 occur. Place frames and frame
mandrels 908 are described in FIGS. 21A-24B.
FIG. 21A is a perspective view of a mandrel with frame plies and
frame mandrels in place in accordance with an embodiment of the
present invention, as shown in FIG. 9. As shown in FIG. 21A,
mandrel 1610 includes recesses for plies, including frame plies
2010, integral flange plies 2040, and wing attachment pocket plies
2050. In one implementation, these frame plies 2010 are placed in
frame recesses 1620 (see FIG. 21B) on mandrel 1610. Once frame
plies 2010 have been placed, frame mandrels 1920 may be placed. In
one implementation, frame mandrels 1920 (as described in FIG. 19)
may be placed upon frame plies 2010 or flange plies 2040 to provide
the support during the cure process. In one implementation, mandrel
1610 may be placed in winding cart (not shown) to allow access for
lay-up. These implementations are merely exemplary, and other
implementations may also be used.
FIG. 21B illustrates frame plies on the mandrel in accordance with
an embodiment of the present invention, as shown in FIG. 21A. As
shown in FIG. 21B, frame plies 2010, may be formed and placed in a
frame recess 1620 on bag 1010, which sits atop media 1210, which
sits atop armature 1020.
FIG. 21C illustrates frame plies and a frame mandrel on the mandrel
in accordance with an embodiment of the present invention, as shown
in FIG. 21A. As shown in FIG. 21C, a frame mandrel, such as frame
mandrel 1920 is placed on top of frame plies 2010 on bag 1010,
which sits atop media 1210, which sits atop armature 1020. This
implementation is merely exemplary, and other implementations may
also be used.
FIGS. 21A-21C have provided an overview of frames and frame
mandrels. FIGS. 22-24B describe frame ply lay-up and frame mandrels
in more detail.
FIG. 22 illustrates wing attachment plies being applied to a
mandrel to form wing attachment pockets in accordance with an
embodiment of the present invention, as shown in FIGS. 21A-21C. As
shown in FIG. 22, wing attachment plies 2050 have been formed and
placed in wing attachment pockets 2210. Wing attachment pockets
2210 may also include metal inserts (not shown). Metal inserts
provide bearing strength in the joint areas.
FIG. 23 illustrates frame plies in frame recesses in a mandrel in
more detail in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C. As shown in FIG. 23, frame
plies 2010 are placed inside frame recess 2310 in mandrel 1410. In
this example, frame plies 2010 are placed inside frame recess 2310
in the forward and lower portion of access door opening 115. Frame
mandrels 1920 may then be placed on frame plies 2010, as shown in
FIG. 24A.
FIGS. 24A-24B, illustrate the combination of frame plies and a
frame mandrel on a mandrel.
FIG. 24A illustrates a frame mandrel in a frame recess in a mandrel
in more detail in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C. As shown in FIG. 24A, frame
mandrel 1920 is placed in frame recess 1620 in mandrel 1610.
FIG. 24B illustrates a frame mandrel over frame plies in a frame
recess in a mandrel in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C, 23, and 24A. As shown in FIG.
24B, frame plies 2010 are placed in frame recess 1620 (not shown).
In this example, frame plies 2010 are placed inside frame recess
1620 in the forward section of access door opening 115. Frame
mandrel 1920 is then placed on top of frame plies 2010. As also
shown in FIG. 24B, the recess in access door 115 may be filled with
a winding aid, such as, door opening filler block 2440. As
described above, winding aids, such as door opening filler block
2440, provide a convex surface for filament winding (described in
the following paragraphs). As shown in FIG. 24B, a second winging
aid 2450 is used to define ply cutting locations during cut and
drape 911 (described below).
FIGS. 21A-24A have described place frames and frame mandrels 908,
as shown in FIG. 9. As shown in FIG. 9, following place frames and
frame mandrels 910, wind inner skin 910 occurs. Wind inner skin 910
is described in FIGS. 25-28.
FIG. 25 illustrates preparing the mandrel for filament winding of
the inner skin in accordance with an embodiment of the present
invention, as shown in FIG. 9. As shown in FIG. 25, mandrel 1390
includes door recess filler 1760, windshield recess filler 1750,
and wing attachment pocket fillers 1740. Other frames and frame
mandrels have also been inserted in mandrel 1390 (as described
above.)
As further shown in FIG. 25, a hoop wrap 2510 may be applied as
hoops 2520 to mandrel 1390 by a filament winding machine (not shown
here, but shown in FIG. 26). Hoop wrap 2510 is wound
circumferentially by the filament winding machine around mandrel
1390, such that space exists between the adjacent hoops 2520. In
one implementation, approximately 4 inches of "advance" is used.
Advance is the space between subsequent winding paths. In the case
of a hoop wrap, advance is the distance between adjacent bands of
fiber being placed by the winding machine head. Hoop wrap 2510 is
subsequently removed from mandrel 1390 after sufficient filament of
desired orientation has been applied to mandrel 1390 to retain
frame plies, frame mandrels, and winding aids.
FIG. 26 illustrates applying filament to the mandrel for filament
winding of the inner skin by a filament winding machine in
accordance with an embodiment of the present invention, as shown in
FIG. 25. As shown in FIG. 26, a filament winding machine 2610
applies filament 2620 (such as carbon fiber) to mandrel 1390. In
the example shown in FIG. 26, approximately 15% of one internal ply
is in place to form the inner skin. In one implementation, when the
inner skin is complete a cross-section of the inner skin will be
about 0.016 inch thick, over about a 0.250 inch thick core. In this
implementation, moreover, frames will generally be about 0.034 inch
thick with a height of about 1.25 inch and a width of about 1.75
inch. These dimensions are provided for exemplary purposes and are
typical of one fuselage structure for one type of aircraft.
Therefore, other implementations may be used, as needed.
As shown in FIG. 26, filament winding machine 2610 is used to apply
an inner skin to mandrel 1390. For filament winding of a structure
(such as mandrel 1390), a filament winding machine having a
capacity of approximately 25 feet in length and a swing of
approximately 3 feet is adequate. An acceptable filament winding
machine for this purpose is commercially available through vendors,
such as Entec in Salt Lake City, Utah. However, other filament
winding machines may be used. In one implementation, the wind angle
may be close to plus or minus 45 degrees, as practicable. For other
implementations, it may be preferable to build a custom winding
machine suited for a particular structure being manufactured. Also,
in some implementations, during filament winding, it may be
necessary to stop and place doubler plies by hand, as needed.
FIG. 27A is a perspective view of a mandrel with a filament-wound
inner skin in accordance with an embodiment of the present
invention, as shown in FIGS. 21A-21C. As shown in FIG. 27A, mandrel
1610 has been fully wound with inner skin 2710 by filament winding
machine 2610 (not shown).
FIG. 27B is a cut-away view of a mandrel with a filament-wound
inner skin in accordance with the embodiment of the present
invention, as shown in FIG. 27A. As shown in FIG. 27B,
filament-wound inner skin 2710 sits atop frame mandrel 1920, which
sits atop frame plies 2010, which sits atop bag 1010, which
surrounds media 1210, which surrounds armature 1020. Filament wound
inner skin 2710 also covers winding aids 2720.
FIG. 28 is a side view of a mandrel with a filament-wound inner
skin with external end hoop plies in accordance with an embodiment
of the present invention, as shown in FIG. 26. As shown in FIG. 28,
external end hoop plies 2810 are placed around portions of mandrel
1390 over inner skin 2820. External hoop plies 2810 are used to
hold inner skin 2810 for cut and drape 911 (described below).
FIGS. 25-28 have described wind inner skin 910, as shown in FIG. 9.
As shown in FIG. 9, following wind inner skin 910, cut and drape
911 occurs. Cut and drape 911 is described in FIGS. 29-30B.
FIG. 29 illustrates cutting a mandrel in accordance with an
embodiment of the present invention, as shown in FIG. 9. As shown
in FIG. 29, ends 2910 have been cut from inner skin 2820 over
mandrel 1390. After ends 2910 have been cut, the other portions of
mandrel 1390 are cut (as described below).
FIG. 30A is a perspective view of a mandrel with inner skin cut and
draped in accordance with an embodiment of the present invention,
as shown in FIG. 27A. As shown in FIG. 30A, mandrel 1610 shows
inner skin 2710. During cut and drape 911 (as described in FIG. 9),
inner skin 2710 is cut in particular locations so that winding aids
can be removed (as described in FIG. 17). For example, as shown in
FIG. 30A, winding aids include door recess fillers and windshield
area fillers. Other winding aids may also include passenger window
recess fillers. Additionally, although not shown in FIG. 30A,
winding aids 2450 may be used around the winding aids to provide a
cutting guide. These winding aids 2450 identify the location of
other winding aids. The winding aids 2450 include raised pins,
which may be used to position cutting aids for removal of excess
material from mandrel 1610.
FIG. 30B is a cut-away view of a mandrel with inner skin that has
been cut and draped in accordance with an embodiment of the
invention, as shown in FIG. 30A. As shown in FIG. 30B, after
cutting inner skin 2710 and removing winding aids in recess areas,
joggle areas 3010 are exposed. Inner skin plies 2710 can now be
draped into joggle areas 3010. Inner skin plies 2710 are draped
into contact with frame plies 2010, frame mandrel 1920, and bag
1010, which surrounds media 1210, which surrounds armature 1020.
Thus, following cutting and removal of winding aids (not shown)
doubler plies 2020 (not shown) are draped to joggle areas 3010.
Doubler plies 2020 are placed over inside corners where skin plies
2710 are cut to allow them to drape into joggled areas. Doubler
plies 2020 reinforce the cut inner skin plies 2710. Joggled areas
3010 are normally located around windshields, windows, and door
openings. Joggled areas 3010 allow for the windows, the windshield,
and the doors to fit flush to the surface of the structure.
Alternatively, joggled areas 3010 could be eliminated, where other
solutions could be used to make the flush fit. These
implementations are merely exemplary, and other implementations may
also be used.
FIGS. 29-30B have described cut and drape 911, as shown in FIG. 9.
As shown in FIG. 9, following cut and drape 911, place core 912
occurs. Machine core to shape 952 and place core 912 are described
in FIGS. 31A-32.
FIG. 31A illustrates machining core in accordance with an
embodiment of the present invention, as shown in FIG. 9. As shown
in FIG. 31A, core details are machined by cutting a desired
peripheral shape from core sheet stock and then chamfering that
periphery to provide sandwich material for placement between inner
and outer skins of the fuselage to enhance skin buckling strength.
Core sheet stock may be prepared to thickness by a core material
supplier or it may be cut to desired thickness in a clean
environment. Core materials include foam core as well as honeycomb
core materials. Foam core materials are made from high temperature
thermoplastics that have been foamed using a blowing agent or some
other foaming methodology. Honeycomb core materials are made from
metal foils or plastic materials (strengthened with natural or
synthetic fibers) formed into paper bonded together in such a
manner as to resemble natural bee's wax honeycomb. Examples of
plastic honeycomb core material include Nomex and Korex materials
registered trademarks of Dupont. However, any type of core material
may be used for machine core to shape 952.
FIG. 31B is a perspective view of a mandrel with core material in
accordance with an embodiment of the present invention, as shown in
FIG. 9. As shown in FIG. 31B, mandrel 1610 includes core pieces
3110, which are applied over film adhesive to the outside of inner
skin 2710, where inner skin 2710 has been draped into recesses in
mandrel 1610. Core pieces 3110 are used to prevent skin buckling.
Core pieces 3110 also help to retain a desired structural
shape.
FIG. 31C is a cut-away view of a mandrel with core details in
accordance with an embodiment of the present invention, as shown in
FIG. 31A. As shown in FIG. 31C, a core piece 3110 is placed in a
recess created in mandrel 1610. Core piece 3110 sits atop inner
skin 2710, which sits atop bag 1010, which sits atop media 1210,
which sits atop armature 1020. Core pieces 3110 can be placed on
inner skin plies 2710 with or without film adhesive. The use of
film adhesive may be needed, if the winding resin being used does
not have adhesive properties.
FIG. 32 illustrates a portion of a mandrel with film adhesive
covering core material in accordance with an embodiment of the
present invention, as shown in FIGS. 31A-31B. As shown in FIG. 32,
core pieces 3110 include core material with film adhesive 3210 and
core material without adhesive (not shown). Separator film 3225 may
also be placed on the inner skin. Separator film 3225, which is
placed to aid manufacture, such as, when it is time to remove
excess material and to drape any joggle areas 3010.
FIGS. 31A-32 have described place core 912, as shown in FIG. 9. As
shown in FIG. 9, following place core 912, wind outer skin 914
occurs. Wind outer skin 914 is described in FIGS. 33-35.
FIG. 33 illustrates preparing a mandrel for application of an outer
skin by a filament winding machine in accordance with an embodiment
of the present invention, as shown in FIG. 9. As shown in FIG. 33,
mandrel 1390 includes end domes 1720, separator film 3225, and end
hoop wraps 2810. End domes 1720 are used to provide a convex
surface. Separator film 3225 is used around the joggle areas to aid
in removal of winding aids in joggle recesses. External end hoop
wraps 2810 are used around mandrel 1390 to hold inner skin plies
2630 in place. Further, as shown in FIG. 33, filament for outer
skin 3330 has begun to be applied to mandrel 1390 by filament
winding machine 2610. Further, hoop wrap 3340 may be applied to
mandrel 1390 by hand or by filament winding machine 2610. Hoop wrap
3340 is wound circumferentially such that gaps exists between
successive wraps but close enough together to hold assorted winding
aids in their correct locations.
FIG. 34 illustrates applying an outer skin to a mandrel by a
filament winding machine in accordance with an embodiment of the
present invention, as shown in FIG. 33. As shown in FIG. 34,
filament winding machine 2610 wraps mandrel 1390 in filament for
outer skin 3410. In this example, filament winding machine 2610 is
at about in a 25% finished state, with outer skin 3410 applied to
mandrel 1390. In one implementation, outer skin 3410 includes two
plies. Alternatively, in another implementation, filament winding
machine 2610 may wind an outer skin 3410 that is twice as thick.
Other implementations may have other plies or other layers.
FIG. 35A is a perspective view of a mandrel with a filament wound
outer skin in accordance with an embodiment of the present
invention, as shown in FIG. 9. As shown in FIG. 35A, filament
winding machine 2610 (not shown) applies outer skin 3510 to mandrel
1610. As also shown in FIG. 35A, filament winding machine 2610 may
apply an outer skin to a large area, which may be larger than just
a fuselage cabin (e.g., to include a tail cone 106).
FIG. 35B is a cut-away view of a mandrel with a filament wound
outer skin in accordance with an embodiment of the present
invention, as shown in FIG. 35A. As shown in FIG. 35B, outer skin
3510 has been placed over core pieces 3110, inner skin 2710, frame
mandrel 1920, and other winding aids. Core piece 3110 sits atop
inner skin 2710, which sits atop bag 1010, which sits atop media
1210, which sits atop armature 1020.
As shown in FIG. 9, following wind outer skin 913, cut and drape
914 occurs. Cut and drape 914 is described in FIGS. 36A-37.
FIG. 36A is a perspective view of a mandrel with outer skin cut and
draped in accordance with an embodiment of the present invention,
as shown in FIG. 9. As shown in FIG. 36A, mandrel 1610 shows outer
skin 3510. During cut and drape 914 (as described in FIG. 9), outer
skin 3510 is cut in particular locations so that winding aids can
be removed. For example, as shown in FIG. 36A, winding aids that
are removed include windshield winding aid 3610, door winding aids
3620, and window winding aid 3630.
FIG. 36B is a cut-away view of a mandrel with outer skin that has
been cut and draped in accordance with an embodiment of the
invention, as shown in FIG. 36A. As shown in FIG. 36B, after
cutting outer skin 3510 and removing winding aids in recess areas
joggle area 3645 is exposed. Outer skin plies 3510 can now be
draped into joggle area 3645. Outer skin plies are draped into
contact with inner skin 2710, which has already been formed into
joggle area 3645. Inner skin plies contact frame plies 2010, frame
mandrel 1920 and bag 1010, which surrounds media 1210, which
surrounds armature 1020. Thus, following cutting and removal of
winding aids (not shown) doubler plies 2020 (not shown) are draped
to joggle area 3645. Doubler plies 2020 are placed over inside
corners where outer skin plies 3510 are cut to allow them to drape
into joggled areas. Doubler plies 2020 reinforce the cut outer skin
plies 3510. Joggled areas 3645 are normally located around
windshields, windows, and door openings. Joggled areas 3645 allow
for the windows, the windshield, and the doors to fit flush to the
surface of the structure. Alternatively, joggled areas 3645 could
be eliminated, where other solutions could be used to make the
flush fit. For example, frames could include joggle areas 3645,
rather than using plies, such as doubler plies 2020.
FIG. 37 illustrates the mandrel after cutting and draping of the
outer skin in accordance with an embodiment of the present
invention, as shown in FIG. 33. As shown in FIG. 37, mandrel 1390
is shown with outer skin 3410 cut and draped into door joggle area
3710. Separator film 3225 (not shown) has been removed from between
inner skin 2630 and outer skin 3410 to allow cutting and draping of
outer skin 3410 into door joggle area 3710. Separator film 3225 is
any low cost thermoplastic film the most prevalent being
polyethylene and nylon. Other films may be used including FEP,
PTFE, and ECTFE. External end hoop plies 2810 are retaining outer
skin 3410 and inner skin 2710 to mandrel 1390. End domes 1720 can
now be removed from both ends of mandrel 1390.
As shown in FIG. 9, following cut and drape 914, close mold 915
occurs. Close mold 915 is described in FIGS. 38A-39D.
FIG. 38A illustrates preparing a circumferential mold for a mandrel
in accordance with an embodiment of the present invention, as shown
in FIG. 9. As shown in FIG. 38, circumferential mold 3810 may
include several pieces (described below). Circumferential mold 3810
may be placed on the exterior of mandrel 1610. In one
implementation, circumferential mold 3810 is approximately 20 feet
long, 4 feet wide, and 6 feet high. In this implementation,
circumferential mold 3810 consists of three pieces: (1) lower
circumferential mold section 3812, (2) left top circumferential
mold section 3814, and (3) right top circumferential mold section
3816. In other implementations, circumferential mold 3810 may be
one piece, two pieces, or more than three pieces. These
implementations are merely exemplary, and other implementations may
also be used.
FIG. 38B is a cut-away view of a mandrel in the circumferential
mold in accordance with an embodiment of the present invention. As
shown in FIG. 38B, circumferential mold 3810 is closed over mandrel
1610. Circumferential mold 3810 covers all portions of mandrel
1610, including outer skin 3510 and outer joggle recesses 3645.
Once circumferential mold 3810 has been closed, a vacuum may be
applied, so that all air is removed between outer skin 3510 and
inner skin 2710, between outer skin 3510 and core piece 2060,
between inner skin 2710 and bag 1010, and between inner skin 2710
and frame mandrel 1920, among other areas. Additionally,
pressurization may be used with a vacuum. In this implementation,
enclosed cavity 1030 containing media 1210 is pressurized. Between
two and three atmospheres are generally adequate for
pressurization, although pressure may vary depending upon the
particular application. If pressurization is used, during cut and
drape 911 and/or cut and drape 914, cuts should be done to allow
for expansion during pressurization. In both of these
implementations, the frame mandrels 1920 may be placed under vacuum
to maintain their shape (described below).
FIG. 39A illustrates preparing a circumferential mold with a vacuum
system for the frame mandrels during curing in accordance with an
embodiment of the present invention, as shown in FIGS. 38A-38B. As
shown in FIG. 39A, vacuum system 3910 may be used so that frame
mandrels (not shown here, but shown in FIG. 39B) maintain the
proper shape. For vacuum system 3910, internal plumbing (not shown)
is needed. Standard vacuum plumbing may accomplish these tasks.
FIG. 39B illustrates a cut-away of the mandrel in the
circumferential mold with a vacuum system for the frame mandrels in
accordance with an embodiment of the present invention, as shown in
FIG. 39A. Vacuum system 3910 provides for pulling a vacuum, using
piping which starts at end plates (not shown), continues through
media 1210, and goes to bag 1010, which then goes to a vacuum port
3920. Thus, a vacuum is transmitted to the frame mandrels, such as
frame mandrel 1920, through vacuum port 3920, which may be
installed when the frame mandrels are positioned in mandrel
1610.
FIG. 39C illustrates a vacuum port in a frame mandrel in accordance
with an embodiment of the present invention, as shown in FIGS. 39A
and 39B. As shown in FIG. 39C, vacuum port 3920 is a couple between
the interior of frame mandrel 1920 and vacuum source tube 3970.
Vacuum from vacuum source tube 3970 is extended into the interior
of frame mandrel 1920, which has media 1210 inside, through
double-ended needle 3925. Double-ended needle 3925 passes through
valve 3965 in frame mandrel 1920 and valve 3965 in mandrel 1610.
Because frame mandrel 1920 is filled with media 1210 it is
necessary to equip frame mandrel with filtering device 3960 to
prevent media 1210 from plugging double ended needle 3925.
FIG. 39D illustrates a device for maintaining a vacuum in a frame
mandrel in accordance with an embodiment of the present invention,
as shown in FIGS. 39B and 39C. As shown in FIG. 39C, vacuum port
3920 includes a double ended inflation needle 3925, of such a
length that one end of needle 3925 when inserted into a valve 3965
in frame mandrel 1920 extends into air space 3962 and the other end
inserted into a valve 3965 in mandrel 1610 extends into vacuum
source tube 3970. Further, needle 3925 is modified at ends 3927,
3928 with side holes 3922 to prevent end plugging and further
comprises a disk 3924 located approximately at the mid-point of
needle 3925 to guard against end plugging by sealing end 3927
against the filtering device 3960.
As shown in FIG. 9, following close mold 915, cure 916 occurs. Cure
916 is described in FIG. 40.
FIG. 40 illustrates curing a filament wound mandrel in a
circumferential mold in an oven in accordance with an embodiment of
the present invention, as shown in FIG. 9. As shown in FIG. 40, in
one implementation, circumferential mold 3810 is placed in oven
4010. In this implementation, oven heat cures the composite
materials on mandrel 1390 against circumferential mold 3810.
Alternatively, heat can be applied using integral heating methods,
such as circulating heated liquid through tubes within
circumferential mold 3810. Alternatively, heat can also be used
inside enclosed cavity 1030 to cure the composite material. Indeed,
any type of oven or any type of heat can be used to cure composite
material inside circumferential mold 3810.
As shown in FIG. 9, following cure 916, de-mold 917 occurs. De-mold
917 is described in FIGS. 41-45. De-mold 917 includes removing
circumferential mold 3810, as described in FIG. 41, removing media
1210, as described in FIG. 42, removing armature 1020, as described
in FIG. 43, removing bag 1010, as described in FIG. 44, and making
bag 1020 available for reuse, as described in FIG. 45.
FIG. 41 illustrates removing a circumferential mold from around a
one-piece integrally stiffened fuselage on a mandrel in accordance
with an embodiment of the present invention, as shown in FIG. 9. As
shown in FIG. 41, fuselage 4110 is removed from circumferential
mold 3810. In one implementation, removal of circumferential mold
3810 occurs after fuselage 4110 has been allowed to cool
sufficiently, for example, to below 150.degree. F. For this
implementation after mandrel 160 has cooled, a vacuum and/or
pressure is also released. Other implementations may be used.
FIG. 42 illustrates removing media from a mandrel in accordance
with an embodiment of the present invention, as shown in FIG. 9. As
shown in FIG. 42, mandrel 1390 is inside one-piece integrally
stiffened fuselage cabin, which in turn is inside circumferential
mold 3810. In one implementation, a vacuum 4210 removes media
through fill ports (not shown) in end plates 1380 (also not shown).
In this implementation, after removal of media from mandrel 1390,
armature 1320 is removed from mandrel 1390 (not shown), and
armature 1320 may then be reused. Other implementations may be also
used.
FIG. 43 illustrates a one-piece integrally stiffened fuselage
contained in a circumferential mold after removal of media and
armature in accordance with one embodiment of the present invention
as shown in FIG. 42. As shown in FIG. 43, media and armature 1320
(not shown) have been removed from mandrel 1390.
FIG. 44 illustrates removing a bag from a one-piece integrally
stiffened fuselage in accordance with an embodiment of the present
invention as shown in FIG. 41. As shown in FIG. 44, bag 1010 is
removed from one-piece integrally stiffened fuselage 4110, while
both are supported on work stands 4410. Bag 1010 as previously
explained has no substantial shape without armature 1020 and media
1210.
FIG. 45 illustrates a bag after removal from a mandrel in
accordance with an embodiment of the present invention, as shown in
FIG. 44. Bag 1010 may now be reused, after removal from one-piece
integrally stiffened fuselage 4110.
As shown in FIG. 9, following de-mold 917, extract frame mandrels
918 occurs. Extract frame mandrels 918 is described in FIG. 46.
FIG. 46 illustrates removing frame mandrels from a one-piece
integrally stiffened fuselage in accordance with an embodiment of
the present invention as shown in FIG. 9. As shown in FIG. 46, once
media has been extracted, and once the bag 1420 has been extracted,
frame mandrels 1920 (not shown) may then be extracted. Once frame
mandrels 1920 have been extracted from mandrel 4200, a structure
can be seen, such as, in this example, one-piece integrally
stiffened fuselage 4610. Fuselage 4610 includes door openings 4610,
windshield opening 4620, and door attachment points 4630. Fuselage
4610 also depicts other components, as shown in FIG. 46.
As shown in FIG. 9, following extract frame mandrels 918, visually
inspect parts 919, trim 920, and prime and paint 921 occurs. One
implementation of actions 919, 920, and 921 is described in FIGS.
47-48. Other implementations may be used.
FIG. 47 illustrates a one-piece integrally-stiffened fuselage
manufactured in accordance with one embodiment of the present
invention as shown in FIG. 9. As shown in FIG. 47, fuselage 4700
has been prepared for inspection. During visually inspecting part
919, fuselage 4700 is examined visually, both interior and exterior
surfaces. In addition, visually inspect part 919 includes
verification that dimensional tolerances are correct. In addition
to visually inspect part 919, trim 920 includes trimming any
material, as needed. Trim 920 includes manual methods (such as a
hand held air powered router motor with router tool) or automatic
methods (such as robot using a router tool). In addition to trim
920, prime and paint 921 includes sanding and filling surfaces to
an acceptable level of smoothness. After sanding and filling,
fuselage 4700 receives paint primer on all exterior surfaces.
FIG. 48 illustrates a one-piece integrally stiffened fuselage
manufactured in accordance with one embodiment of the present
invention as shown in FIG. 9. As shown in FIG. 48, fuselage 4800
includes both a fuselage cabin 4810 and a tail cone 4820. FIG. 48
demonstrates that fuselage 4800 includes both fuselage cabin 4810
and tail cone 4820. In other implementations, other parts of an
aircraft may further be included with fuselage 4800.
As shown in FIG. 9, following prime and paint 921, store for
assembly 922 occurs. Store for assembly involves storing a
structure, such as a fuselage, until needed, e.g., until the
fuselage is needed to assemble an aircraft.
As described in the preceding sections, various implementations may
be used to create a structure, such as fuselage 4700 or fuselage
4800. The following section illustrates one of many such
alternative implementations.
2. Alternate 2
FIG. 49 is a block diagram illustrating the process of
manufacturing a one-piece fuselage in accordance with another
embodiment of the present invention, as shown in FIG. 9. As shown
in FIG. 49, system 4900 is substantially similar to system 900
shown in FIG. 9. However, as shown in FIG. 49, close composite mold
4906 has replaced close mold 915, prepare autoclave 4904 has
replaced prepare oven 975, and cure using autoclave 4902 has
replaced cure 916. Close composite mold 4906 involves the use of a
mold manufactured from composite materials. In one implementation,
the mold is manufactured from high temperature fiber-reinforced
plastics. However, in other implementations, other molds and other
materials could be used. Prepare autoclave 4904 and cure using
autoclave show that an autoclave is used to cure the structure in
system 4900. Cure using autoclave pressure is described below.
Close composite mold 4906, prepare autoclave 4904, and cure using
autoclave 4902 are described in FIGS. 50-53.
FIG. 50 illustrates assembling a circumferential mold around a
mandrel in accordance with an embodiment of the present invention,
as shown in FIG. 49. As shown in FIG. 50, fuselage 5000 may be
prepared for curing according to prepare autoclave 4904 (from FIG.
49). Fuselage 5000 may be placed in a mold 5010 for curing. In one
implementation, mold 5010 consists of seven pieces. Four of the
seven pieces are visible in FIG. 50: (1) upper left half 5002, (2)
upper right half 5004, (3) lower forward segment 5008, and (4)
lower aft segment 5006. The other segments (not shown) comprise:
(1) left windshield area 5012 (not shown), (2) the right windshield
area 5014 (not shown), and (3) the bulkhead flange upper half
segment 5016 (not shown). In other implementations, more or less
pieces could be used to construct the mold These implementations
are merely exemplary, and other implementations may also be
used.
FIG. 51 illustrates bagging a circumferential mold in accordance
with an embodiment of the present invention, as shown in FIG. 50.
As shown in FIG. 51, the seven pieces of mold 5010 (as described in
FIG. 50) have been placed around fuselage 5000. After construction
of the mold 5010, it is placed in a bag 5102 for sealing to form
bag assembly 5100. Bag 5102 may be made of nylon. However, bag 5102
may also be made of any other material. In one implementation, bag
5102 is sealed such that a vacuum is formed between the bag 1420
and envelope bag 5102. In this implementation, the vacuum provides
for a void-free structure. Other implementations may also be
used.
FIG. 52 illustrates placing a circumferential mold in an autoclave
for curing in accordance with an embodiment of the present
invention, as shown in FIG. 51. As shown in FIG. 52, bag assembly
5200 may be placed in an autoclave 5202. In one implementation,
autoclave 5202 applies pressure to bag assembly 5200. In this
implementation, during autoclave curing, the frames and frame
mandrels are maintained under by vacuum to maintain the proper
shape for the frames. In this implementation, further curing by
autoclave 5202 generally takes 11/2 to 2 hours at a temperature
between 250.degree. F. and 350.degree. F. at one to three
atmospheres of pressure. The pressure is applied to the outside of
bag 5102 and to the media cavity between armature 1320 and bag
1420. Autoclaves such as those manufactured by Thermal Equipment
Corporation, Taricco Corporation, McGill AirPressure Corporation,
Melco Steel Incorporated, and American Autoclave Company may be
used. These implementations are only exemplary, and other
implementations and other types of autoclaves may also be used.
FIG. 53 illustrates removing a circumferential mold after curing in
an autoclave in accordance with an embodiment of the present
invention, as shown in FIG. 52. As shown in FIG. 53, the pieces of
mold 5010 are removed following curing by autoclave 5202. In FIG.
53, upper left half 5002 is removed, while upper right half 5004 is
still in place.
As shown in FIG. 49, following de-mold 917, extract mandrels 918,
visually inspect part 919, trim 920, and prime and paint 921
occurs. One implementation of actions 919, 920, and 921 is depicted
in FIG. 54. Other implementations may be used.
FIG. 54 illustrates a one-piece integrally-stiffened fuselage
manufactured in accordance with another embodiment of the present
invention, as shown in FIG. 49. As shown in FIG. 54, fuselage 5400
has been prepared for inspection. During visually inspect parts
919, fuselage 5400 is examined visually, both interior and exterior
surfaces. In addition, visually inspect parts 919 includes
verification that dimensional tolerances are correct. In addition
to visually inspect parts 919, trim 920 includes trimming any
material, as needed. Trim 920 includes manual methods (such as a
hand held air powered router motor with router tool) or automatic
methods (such as robot using a router tool). In addition to trim
920, prime and paint 921 includes sanding and filling surfaces to
an acceptable level of smoothness. After sanding and filling,
fuselage 5400 receives paint primer on all exterior surfaces.
3. Other Alternates:
Alternates 1 (such as process alternate #1805) and Alternate 2
(such as process alternate #2815) are described herein, but any
number of alternate methods and structures are possible for a
one-piece structure, such as a fuselage, using the claimed
invention.
VI. CONCLUSION
As described above, therefore, other embodiments of the invention
will be apparent to those skilled in the art from consideration of
the specification and practice of the invention disclosed herein.
It is intended that the specification and examples be considered as
exemplary only, with a true scope and spirit of the invention being
indicated by the following claims and their equivalents. In this
context, equivalents mean each and every implementation for
carrying out the functions recited in the claims, even if not
explicitly described therein.
* * * * *